JP6550230B2 - Element - Google Patents

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JP6550230B2
JP6550230B2 JP2014245346A JP2014245346A JP6550230B2 JP 6550230 B2 JP6550230 B2 JP 6550230B2 JP 2014245346 A JP2014245346 A JP 2014245346A JP 2014245346 A JP2014245346 A JP 2014245346A JP 6550230 B2 JP6550230 B2 JP 6550230B2
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carbon fiber
composite member
metal
aircraft
heat
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JP2016108398A5 (en
JP2016108398A (en
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信幸 神原
信幸 神原
阿部　俊夫
俊夫 阿部
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Mitsubishi Heavy Industries Ltd
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Mitsubishi Heavy Industries Ltd
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Priority to JP2014245346A priority Critical patent/JP6550230B2/en
Priority to EP15864708.1A priority patent/EP3214112B1/en
Priority to PCT/JP2015/080002 priority patent/WO2016088470A1/en
Priority to US15/531,606 priority patent/US20180281983A1/en
Publication of JP2016108398A publication Critical patent/JP2016108398A/en
Publication of JP2016108398A5 publication Critical patent/JP2016108398A5/ja
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F28—HEAT EXCHANGE IN GENERAL
    • F28F—DETAILS OF HEAT-EXCHANGE AND HEAT-TRANSFER APPARATUS, OF GENERAL APPLICATION
    • F28F21/00—Constructions of heat-exchange apparatus characterised by the selection of particular materials
    • F28F21/06—Constructions of heat-exchange apparatus characterised by the selection of particular materials of plastics material
    • F28F21/067—Details
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B5/00—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts
    • B32B5/02—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by structural features of a fibrous or filamentary layer
    • B32B5/12—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by structural features of a fibrous or filamentary layer characterised by the relative arrangement of fibres or filaments of different layers, e.g. the fibres or filaments being parallel or perpendicular to each other
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B5/00—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts
    • B32B5/22—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by the presence of two or more layers which are next to each other and are fibrous, filamentary, formed of particles or foamed
    • B32B5/24—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by the presence of two or more layers which are next to each other and are fibrous, filamentary, formed of particles or foamed one layer being a fibrous or filamentary layer
    • B32B5/26—Layered products characterised by the non- homogeneity or physical structure, i.e. comprising a fibrous, filamentary, particulate or foam layer; Layered products characterised by having a layer differing constitutionally or physically in different parts characterised by the presence of two or more layers which are next to each other and are fibrous, filamentary, formed of particles or foamed one layer being a fibrous or filamentary layer another layer next to it also being fibrous or filamentary
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08J—WORKING-UP; GENERAL PROCESSES OF COMPOUNDING; AFTER-TREATMENT NOT COVERED BY SUBCLASSES C08B, C08C, C08F, C08G or C08H
    • C08J5/00—Manufacture of articles or shaped materials containing macromolecular substances
    • C08J5/04—Reinforcing macromolecular compounds with loose or coherent fibrous material
    • C08J5/0405—Reinforcing macromolecular compounds with loose or coherent fibrous material with inorganic fibres
    • C08J5/042—Reinforcing macromolecular compounds with loose or coherent fibrous material with inorganic fibres with carbon fibres
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08J—WORKING-UP; GENERAL PROCESSES OF COMPOUNDING; AFTER-TREATMENT NOT COVERED BY SUBCLASSES C08B, C08C, C08F, C08G or C08H
    • C08J5/00—Manufacture of articles or shaped materials containing macromolecular substances
    • C08J5/04—Reinforcing macromolecular compounds with loose or coherent fibrous material
    • C08J5/047—Reinforcing macromolecular compounds with loose or coherent fibrous material with mixed fibrous material
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08J—WORKING-UP; GENERAL PROCESSES OF COMPOUNDING; AFTER-TREATMENT NOT COVERED BY SUBCLASSES C08B, C08C, C08F, C08G or C08H
    • C08J5/00—Manufacture of articles or shaped materials containing macromolecular substances
    • C08J5/04—Reinforcing macromolecular compounds with loose or coherent fibrous material
    • C08J5/06—Reinforcing macromolecular compounds with loose or coherent fibrous material using pretreated fibrous materials
    • C—CHEMISTRY; METALLURGY
    • C09—DYES; PAINTS; POLISHES; NATURAL RESINS; ADHESIVES; COMPOSITIONS NOT OTHERWISE PROVIDED FOR; APPLICATIONS OF MATERIALS NOT OTHERWISE PROVIDED FOR
    • C09K—MATERIALS FOR MISCELLANEOUS APPLICATIONS, NOT PROVIDED FOR ELSEWHERE
    • C09K5/00—Heat-transfer, heat-exchange or heat-storage materials, e.g. refrigerants; Materials for the production of heat or cold by chemical reactions other than by combustion
    • C09K5/08—Materials not undergoing a change of physical state when used
    • C09K5/14—Solid materials, e.g. powdery or granular
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B2262/00—Composition or structural features of fibres which form a fibrous or filamentary layer or are present as additives
    • B32B2262/10—Inorganic fibres
    • B32B2262/106—Carbon fibres, e.g. graphite fibres
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B32—LAYERED PRODUCTS
    • B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
    • B32B2605/00—Vehicles
    • B32B2605/18—Aircraft
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/40—Sound or heat insulation, e.g. using insulation blankets
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C2001/0054—Fuselage structures substantially made from particular materials
    • B64C2001/0072—Fuselage structures substantially made from particular materials from composite materials
    • C—CHEMISTRY; METALLURGY
    • C08—ORGANIC MACROMOLECULAR COMPOUNDS; THEIR PREPARATION OR CHEMICAL WORKING-UP; COMPOSITIONS BASED THEREON
    • C08J—WORKING-UP; GENERAL PROCESSES OF COMPOUNDING; AFTER-TREATMENT NOT COVERED BY SUBCLASSES C08B, C08C, C08F, C08G or C08H
    • C08J2363/00—Characterised by the use of epoxy resins; Derivatives of epoxy resins
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F28—HEAT EXCHANGE IN GENERAL
    • F28F—DETAILS OF HEAT-EXCHANGE AND HEAT-TRANSFER APPARATUS, OF GENERAL APPLICATION
    • F28F13/00—Arrangements for modifying heat-transfer, e.g. increasing, decreasing
    • F28F2013/001—Particular heat conductive materials, e.g. superconductive elements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F28—HEAT EXCHANGE IN GENERAL
    • F28F—DETAILS OF HEAT-EXCHANGE AND HEAT-TRANSFER APPARATUS, OF GENERAL APPLICATION
    • F28F2255/00—Heat exchanger elements made of materials having special features or resulting from particular manufacturing processes
    • F28F2255/06—Heat exchanger elements made of materials having special features or resulting from particular manufacturing processes composite, e.g. polymers with fillers or fibres
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/40—Weight reduction

Description

本発明は、航空機又は人工衛星などに使用される部材に関する。   The present invention relates to a member used for an aircraft or a satellite.

航空機は、電子機器、バッテリ、及びエンジンのような発熱部を有する。発熱部で発生した熱は、熱流路材と呼ばれる部材を介して放出（排熱）される。従来、航空機の熱流路材として、金属シート又はジャンパー線が用いられている。人工衛星において使用される紙製のハニカムコアを有する熱流制御体の一例が特許文献１に開示されている。   An aircraft has heat generating components such as electronics, batteries, and engines. The heat generated in the heat generating portion is released (exhaust heat) through a member called a heat flow passage material. Conventionally, a metal sheet or a jumper wire is used as a thermal flow path material of an aircraft. Patent Document 1 discloses an example of a heat flow control body having a paper honeycomb core used in a satellite.

実開平０５−０８６７９９号公報Japanese Utility Model Application Publication No. 05-086799

近年において、航空機の部材の材料は、金属から複合材に遷移している。また、航空機の高度化に伴い、電子機器の使用が増大している。すなわち、近年においては、発熱部の発熱量が増大しているのに対し、航空機の部材の熱伝導率は低下している。そのため、発熱部で発生した熱を効率良く放出できる技術の案出が要望される。   In recent years, the materials of aircraft components have transitioned from metal to composites. Also, with the advancement of aircraft, the use of electronic devices is increasing. That is, in recent years, while the calorific value of the heat generating portion has increased, the thermal conductivity of the members of the aircraft has decreased. Therefore, it is required to devise a technology capable of efficiently releasing the heat generated in the heat generating portion.

本発明の態様は、航空機又は人工衛星などの発熱部で発生した熱を効率良く放出できる部材を提供することを目的とする。   An aspect of the present invention is to provide a member capable of efficiently releasing heat generated in a heat generating portion such as an aircraft or a satellite.

本発明の第１の態様に従えば、金属コート炭素繊維及びピッチ系炭素繊維の一方又は両方を含む熱伝導性炭素繊維で強化されたプラスチックを有する第１複合部材を備え、繊維方向に関して前記熱伝導性炭素繊維の一端部が発熱部に配置され、前記熱伝導性炭素繊維の他端部が放熱部に配置される部材が提供される。   According to a first aspect of the present invention, there is provided a first composite member comprising a thermally conductive carbon fiber reinforced plastic comprising metal coated carbon fiber and / or pitch based carbon fiber, said heat in relation to the fiber direction A member is provided, wherein one end of the conductive carbon fiber is disposed in the heat generating portion, and the other end of the thermally conductive carbon fiber is disposed in the heat radiating portion.

本発明の第１の態様によれば、第１複合部材が金属コート炭素繊維及びピッチ系炭素繊維の一方又は両方を含む熱伝導性炭素繊維で強化された炭素繊維強化プラスチックを含み、その熱伝導性炭素繊維の一端部が発熱部に配置され、他端部が放熱部に配置される。そのため、発熱部で発生した熱は、熱伝導性炭素繊維を伝わって、効率良く放熱部に放出される。また、第１複合部材は、航空機又は人工衛星などの強度部材として使用可能である。そのため、金属シート又はジャンパー線のような専用の熱流路材を別途設けなくても、発熱部の熱を放熱部に放出することができる。そのため、重量の増大を抑制しつつ、発熱部で発生した熱が、効率良く放熱部に放出される。また、熱伝導性炭素繊維とプラスチックとは一緒に成型されるので、ロバスト性が向上する。   According to a first aspect of the present invention, a first composite member comprises a carbon fiber reinforced plastic reinforced with a thermally conductive carbon fiber comprising one or both of metal coated carbon fiber and pitch carbon fiber, the thermal conductivity of which One end of the carbon fiber is disposed in the heat generating portion, and the other end is disposed in the heat radiating portion. Therefore, the heat generated in the heat generating portion is conducted to the heat conductive carbon fiber and is efficiently released to the heat radiating portion. In addition, the first composite member can be used as a strength member such as an aircraft or a satellite. Therefore, the heat of the heat generating portion can be released to the heat radiating portion without separately providing a dedicated heat flow passage material such as a metal sheet or a jumper wire. Therefore, the heat generated in the heat generating portion is efficiently released to the heat radiating portion while suppressing an increase in weight. Also, since the thermally conductive carbon fiber and the plastic are molded together, the robustness is improved.

本発明の第２の態様に従えば、金属コート炭素繊維及びピッチ系炭素繊維の一方又は両方を含む熱伝導性炭素繊維で強化されたプラスチックを有する第１複合部材を備え、繊維方向に関して前記熱伝導性炭素繊維の中央部が発熱部に配置され、前記熱伝導性炭素繊維の一端部及び他端部が放熱部に配置される部材が提供される。   According to a second aspect of the present invention, there is provided a first composite member comprising a thermally conductive carbon fiber reinforced plastic comprising metal coated carbon fibers and / or pitch based carbon fibers, said heat in relation to the fiber direction A member is provided, wherein a central portion of the conductive carbon fiber is disposed in the heat generating portion, and one end and the other end of the thermally conductive carbon fiber are disposed in the heat radiating portion.

本発明の第２の態様によれば、熱伝導性炭素繊維の中央部が発熱部に配置され、一端部及び他端部が放熱部に配置されるので、発熱部で発生した熱は、熱伝導性炭素繊維を伝わって、効率良く放熱部に放出される。そのため、重量の増大を抑制しつつ、発熱部で発生した熱が、効率良く放熱部に放出される。また、熱伝導性炭素繊維とプラスチックとは一緒に成型されるので、ロバスト性が向上する。   According to the second aspect of the present invention, the central portion of the thermally conductive carbon fiber is disposed in the heat generating portion, and the one end portion and the other end are disposed in the heat radiating portion. It travels through the conductive carbon fiber and is efficiently released to the heat sink. Therefore, the heat generated in the heat generating portion is efficiently released to the heat radiating portion while suppressing an increase in weight. Also, since the thermally conductive carbon fiber and the plastic are molded together, the robustness is improved.

本発明の第１の態様又は第２の態様において、前記第１複合部材は、前記繊維方向と交差する並列方向に配置された複数の前記熱伝導性炭素繊維を含むプリプレグシートを前記繊維方向及び前記並列方向と交差する積層方向に複数積層した積層体を含み、前記繊維方向に関する熱伝導率は、前記並列方向に関する熱伝導率及び前記積層方向に関する熱伝導率よりも大きくてもよい。   In the first aspect or the second aspect of the present invention, the first composite member includes a prepreg sheet including a plurality of the thermally conductive carbon fibers arranged in a parallel direction crossing the fiber direction, and the fiber direction and The heat conductivity with respect to the fiber direction may be greater than the heat conductivity with respect to the parallel direction and the heat conductivity with respect to the stacking direction.

これにより、熱伝導率に異方性が付与されるので、発熱部で発生した熱が、並列方向及び積層方向に伝わることが抑制され、放熱部に効率良く放出される。   Thereby, the thermal conductivity is imparted with anisotropy, so that the heat generated in the heat generating portion is suppressed from being transmitted in the parallel direction and the stacking direction, and is efficiently released to the heat radiating portion.

本発明の第１の態様及び第２の態様において、前記第１複合部材は板状部材であり、前記第１複合部材の表面及び裏面の一方又は両方に配置され、炭素繊維で強化されたプラスチックを含む第２複合部材を備え、前記熱伝導性炭素繊維の一端部及び他端部のそれぞれは露出してもよい。   In the first aspect and the second aspect of the present invention, the first composite member is a plate-like member, and a carbon fiber reinforced plastic is disposed on one or both of the front surface and the back surface of the first composite member. And the other end of the thermally conductive carbon fiber may be exposed.

これにより、第１複合部材が第２複合部材で支持され、強度が維持される。熱伝導性炭素繊維の一端部及び他端部のそれぞれは、第２複合部材で覆われずに露出するので、発熱部で発生した熱は、放熱部から効率良く放出される。   Thereby, the first composite member is supported by the second composite member, and the strength is maintained. Each of the one end portion and the other end portion of the thermally conductive carbon fiber is exposed without being covered by the second composite member, so the heat generated in the heat generating portion is efficiently released from the heat radiating portion.

本発明の第１の態様及び第２の態様において、前記発熱部は、前記航空機の電子機器を含み、前記放熱部は、前記航空機の燃料タンクを含んでもよい。   In the first aspect and the second aspect of the present invention, the heat generating portion may include the electronic device of the aircraft, and the heat radiating portion may include a fuel tank of the aircraft.

これにより、電子機器の発熱量が増大しても、電子機器で発生した熱は、燃料タンクに効率良く放出される。   Thereby, even if the calorific value of the electronic device increases, the heat generated by the electronic device is efficiently released to the fuel tank.

本発明の態様によれば、発熱部で発生した熱を効率良く放出できる部材が提供される。   According to an aspect of the present invention, a member capable of efficiently releasing the heat generated in the heat generating portion is provided.

図１は、第１実施形態に係る航空機の一例を示す図である。FIG. 1 is a view showing an example of an aircraft according to the first embodiment. 図２は、第１実施形態に係る複合部材の一例を模式的に示す斜視図である。FIG. 2 is a perspective view schematically showing an example of the composite member according to the first embodiment. 図３は、第１実施形態に係る複合部材の製造方法の一例を模式的に示す図である。FIG. 3: is a figure which shows typically an example of the manufacturing method of the composite member which concerns on 1st Embodiment. 図４は、第１実施形態に係る部材の一例を示す断面図である。FIG. 4 is a cross-sectional view showing an example of a member according to the first embodiment. 図５は、第２実施形態に係る部材の一例を示す断面図である。FIG. 5 is a cross-sectional view showing an example of a member according to the second embodiment. 図６は、第３実施形態に係る航空機の主翼の一例を示す平面図である。FIG. 6 is a plan view showing an example of the main wing of the aircraft according to the third embodiment. 図７は、第３実施形態に係るシアタイの一例を概略的に示す図である。FIG. 7 is a view schematically showing an example of a shear tie according to a third embodiment. 図８は、第３実施形態に係るシアタイの一例を示す断面図である。FIG. 8 is a cross-sectional view showing an example of a shear tie according to a third embodiment.

以下、本発明に係る実施形態について図面を参照しながら説明するが、本発明はこれに限定されない。以下で説明する各実施形態の構成要素は、適宜組み合わせることができる。また、一部の構成要素を用いない場合もある。   Hereinafter, embodiments of the present invention will be described with reference to the drawings, but the present invention is not limited thereto. The components of each embodiment described below can be combined as appropriate. In addition, some components may not be used.

＜第１実施形態＞
第１実施形態について説明する。図１は、本実施形態に係る航空機１の一例を示す図である。図１に示すように、航空機１は、胴体２と、主翼３と、水平尾翼４と、垂直尾翼５と、エンジン６と、燃料タンク７と、コックピット８と、電子機器９と、バッテリ１０と、を備えている。 First Embodiment
The first embodiment will be described. FIG. 1 is a view showing an example of an aircraft 1 according to the present embodiment. As shown in FIG. 1, the aircraft 1 includes a fuselage 2, a main wing 3, a horizontal tail 4, a vertical tail 5, an engine 6, a fuel tank 7, a cockpit 8, an electronic device 9, and a battery 10. And.

胴体２、主翼３、水平尾翼４、及び垂直尾翼５の少なくとも一部は、複合材で形成されている。複合材は、炭素繊維で強化されたプラスチックである炭素繊維強化プラスチック（carbon fiber reinforced plastic：ＣＦＲＰ）を含む。なお、複合材が、ガラス繊維で強化されたプラスチックであるガラス繊維強化プラスチック（glass fiber reinforced plastic：ＧＦＲＰ）を含んでもよい。   At least a portion of the fuselage 2, the main wing 3, the horizontal tail 4 and the vertical tail 5 is formed of a composite material. The composite includes carbon fiber reinforced plastic (CFRP), which is a carbon fiber reinforced plastic. The composite material may include glass fiber reinforced plastic (GFRP) which is a glass fiber reinforced plastic.

なお、胴体２、主翼３、水平尾翼４、及び垂直尾翼５の少なくとも一部が、アルミニウム合金（ジュラルミン）のような金属で形成されてもよい。   Note that at least a portion of the fuselage 2, the main wing 3, the horizontal tail wing 4, and the vertical tail wing 5 may be formed of a metal such as an aluminum alloy (duralmin).

本実施形態において、航空機１の部材の少なくとも一部は、金属コート炭素繊維（ＭＣ）で強化されたプラスチックである金属コート炭素繊維強化プラスチックを有する。金属コート炭素繊維で強化されたプラスチックは、ＭＣ−ＣＦＲＰ、とも呼ばれる。金属コート炭素繊維は、熱伝導性を有する熱伝導性炭素繊維である。   In the present embodiment, at least a part of the members of the aircraft 1 has a metal-coated carbon fiber reinforced plastic which is a metal-reinforced carbon fiber (MC) reinforced plastic. Plastics reinforced with metal coated carbon fibers are also called MC-CFRP. The metal coated carbon fiber is a thermally conductive carbon fiber having thermal conductivity.

図２は、本実施形態に係る金属コート炭素繊維強化プラスチックを含む複合部材１１の一例を模式的に示す斜視図である。図２に示すように、複合部材１１は、金属コート炭素繊維１２で強化されたプラスチック１３である金属コート炭素繊維強化プラスチック１４を有する。金属コート炭素繊維１２は、炭素繊維１５と、炭素繊維１５に被覆された金属１６とを有する。   FIG. 2: is a perspective view which shows typically an example of the composite member 11 containing the metal coat carbon fiber reinforced plastic which concerns on this embodiment. As shown in FIG. 2, the composite member 11 has a metal-coated carbon fiber reinforced plastic 14 which is a plastic 13 reinforced with metal-coated carbon fibers 12. The metal-coated carbon fiber 12 has a carbon fiber 15 and a metal 16 coated on the carbon fiber 15.

複合部材１１は、複数の金属コート炭素繊維１２を有する。金属コート炭素繊維１２は、第１の方向に長い。複数の金属コート炭素繊維１２が、第１の方向と直交する第２の方向に並列に並べられる。また、複数の金属コート炭素繊維１２が、第１の方向及び第２の方向と直交する第３の方向に配置される。   The composite member 11 has a plurality of metal coated carbon fibers 12. The metal coated carbon fiber 12 is long in the first direction. A plurality of metal coated carbon fibers 12 are arranged in parallel in a second direction orthogonal to the first direction. Also, a plurality of metal coated carbon fibers 12 are disposed in a third direction orthogonal to the first direction and the second direction.

以下の説明においては、金属コート炭素繊維１２の長手方向（第１の方向）を適宜、繊維方向、と称する。また、以下の説明においては、複数の金属コート炭素繊維１２が並べられる方向（第２の方向）を適宜、並列方向、と称する。また、以下の説明においては、複数の金属コート炭素繊維１２が配置される方向（第３の方向）を適宜、積層方向、と称する。   In the following description, the longitudinal direction (first direction) of the metal-coated carbon fiber 12 is appropriately referred to as a fiber direction. Moreover, in the following description, the direction (second direction) in which the plurality of metal-coated carbon fibers 12 are arranged is appropriately referred to as a parallel direction. Moreover, in the following description, the direction (third direction) in which the plurality of metal-coated carbon fibers 12 are arranged is appropriately referred to as a lamination direction.

複数の金属コート炭素繊維１２は、並列方向及び積層方向のそれぞれに関して、間隔をあけて配置される。複数の金属コート炭素繊維１２の間に、プラスチック１３が配置される。本実施形態において、プラスチック１３は、エポキシ樹脂を含む。   The plurality of metal-coated carbon fibers 12 are spaced apart in each of the parallel direction and the stacking direction. The plastic 13 is disposed between the plurality of metal coated carbon fibers 12. In the present embodiment, the plastic 13 contains an epoxy resin.

金属コート炭素繊維１２の炭素繊維１５の直径は、例えば、５μｍ以上１０μｍ以下である。炭素繊維１５の表面に金属１６が被覆されている。金属１６の熱伝導率は、炭素繊維１５の熱伝導率よりも高い。プラスチック１３の熱伝導率は、金属１６の熱伝導率及び炭素繊維１５の熱伝導率よりも低い。すなわち、プラスチック１３の熱伝導率は、金属コート炭素繊維１２の熱伝導率よりも低い。   The diameter of the carbon fiber 15 of the metal-coated carbon fiber 12 is, for example, 5 μm or more and 10 μm or less. The metal 16 is coated on the surface of the carbon fiber 15. The thermal conductivity of the metal 16 is higher than the thermal conductivity of the carbon fiber 15. The thermal conductivity of the plastic 13 is lower than the thermal conductivity of the metal 16 and the thermal conductivity of the carbon fiber 15. That is, the thermal conductivity of the plastic 13 is lower than the thermal conductivity of the metal-coated carbon fiber 12.

本実施形態において、金属１６は、ニッケルである。金属コート炭素繊維１２は、ニッケルコート炭素繊維である。なお、金属１６は、金、銀、及び銅の少なくとも一つでもよい。   In the present embodiment, the metal 16 is nickel. The metal coated carbon fiber 12 is a nickel coated carbon fiber. The metal 16 may be at least one of gold, silver and copper.

図３は、本実施形態に係る複合部材１１の製造方法の一例を模式的に示す図である。図３の（ステップＡ）に示すように、金属コート炭素繊維１２が製造される。金属コート炭素繊維１２は、直径が５μｍから１０μｍ程度の炭素繊維１５に金属１６を被覆することによって製造される。本実施形態においては、金属１６として、ニッケルが被覆される。なお、金属１６として、金、銀、及び銅の少なくとも一つが被覆されてもよい。   FIG. 3: is a figure which shows typically an example of the manufacturing method of the composite member 11 which concerns on this embodiment. As shown in FIG. 3 (Step A), metal coated carbon fibers 12 are produced. The metal-coated carbon fiber 12 is manufactured by coating a metal 16 on a carbon fiber 15 having a diameter of about 5 μm to 10 μm. In the present embodiment, nickel is coated as the metal 16. Note that at least one of gold, silver, and copper may be coated as the metal 16.

図３の（ステップＢ）に示すように、複数の金属コート炭素繊維１２が並列方向に並べられ、エポキシ樹脂のようなブラスチック１３で固められる。複数の金属コート炭素繊維１２は、撚らずに、引き揃えられた状態で、プラスチック１３で固められる。   As shown in FIG. 3 (Step B), a plurality of metal-coated carbon fibers 12 are arranged in a parallel direction and hardened with a plastic 13 such as an epoxy resin. The plurality of metal coated carbon fibers 12 are consolidated with the plastic 13 in a straightened state without twisting.

並列方向は、繊維方向に配置された複数の金属コート炭素繊維１２が並べられる方向である。繊維方向と並列方向とは直交する。   The parallel direction is a direction in which the plurality of metal-coated carbon fibers 12 arranged in the fiber direction are arranged. The fiber direction and the parallel direction are orthogonal to each other.

プラスチック１３と、並列方向に配置されプラスチック１３で固められた複数の金属コート炭素繊維１２とを含むシート状の部材は、プリプレグシート１７、と呼ばれる。   A sheet-like member including a plastic 13 and a plurality of metal-coated carbon fibers 12 arranged in a parallel direction and hardened with the plastic 13 is called a prepreg sheet 17.

図３の（ステップＣ）に示すように、製造された複数のプリプレグシート１７が積層方向に積層される。   As shown in (Step C) of FIG. 3, a plurality of manufactured prepreg sheets 17 are stacked in the stacking direction.

積層方向は、プリプレグシート１７が複数積層される方向である。積層方向は、繊維方向及び並列方向と直交する。   The stacking direction is a direction in which a plurality of prepreg sheets 17 are stacked. The stacking direction is orthogonal to the fiber direction and the parallel direction.

プリプレグシート１７の積層体は、オートクレーブと呼ばれる加熱加圧装置で、高温高圧下で加熱処理される。これにより、複数のプリプレグシート１７の積層体である複合部材１１が製造される。   The laminate of the prepreg sheet 17 is heat-treated under high temperature and high pressure by a heating and pressurizing device called an autoclave. Thereby, the composite member 11 which is a laminate of the plurality of prepreg sheets 17 is manufactured.

なお、本実施形態においては、複数のプリプレグシート１７の金属コート炭素繊維１２が全て同一方向に配置される、所謂、一方向積層であることとした。第１のプリプレグシート１７の金属コート繊維１２が第１の方向に配置され、第１のプリプレグシート１７に重なる第２のプリプレグシート１７の金属コート繊維１２が第１のプリプレグシート１７の第１の方向と交差する第２の方向に配置される、所謂、クロスプライ積層でもよい。   In the present embodiment, all the metal-coated carbon fibers 12 of the plurality of prepreg sheets 17 are arranged in the same direction, that is, so-called unidirectional lamination. The metal-coated fibers 12 of the first prepreg sheet 17 are disposed in the first direction, and the metal-coated fibers 12 of the second prepreg sheet 17 overlapping the first prepreg sheet 17 are the first of the first prepreg sheets 17. It may be a so-called cross-ply laminate disposed in a second direction intersecting the direction.

図４は、本実施形態に係る航空機１用の部材（熱流路材）２０の一例を示す断面図である。本実施形態において、部材２０は、複合部材１１と、複合部材１１に接続される複合部材２１と、を含む。部材２０は、胴体２、主翼３、水平尾翼４、及び垂直尾翼５の少なくとも一部に使用される。   FIG. 4 is a cross-sectional view showing an example of a member (heat flow path material) 20 for the aircraft 1 according to the present embodiment. In the present embodiment, the member 20 includes the composite member 11 and the composite member 21 connected to the composite member 11. The member 20 is used for at least a part of the fuselage 2, the main wing 3, the horizontal tail 4 and the vertical tail 5.

例えば、金属コート炭素繊維１２を含むプリプレグシートと、金属が被覆されていない炭素繊維を含むプリブレクシートとが積層され、その積層体がオートクレーブで加熱加圧処理されることによって、部材２０が製造されてもよい。   For example, a member 20 is manufactured by laminating a prepreg sheet containing metal coated carbon fibers 12 and a pre-bleaked sheet containing carbon fibers not coated with a metal, and subjecting the laminated body to heating and pressing in an autoclave. May be

図４に示すように、繊維方向に関して金属コート炭素繊維１２の一端部１２Ａが航空機１の発熱部に配置される。繊維方向に関して金属コート炭素繊維１２の他端部１２Ｂが航空機１の放熱部に配置される。   As shown in FIG. 4, one end 12 A of the metal-coated carbon fiber 12 is disposed in the heat generating portion of the aircraft 1 in the fiber direction. The other end 12 B of the metal-coated carbon fiber 12 in the fiber direction is disposed in the heat dissipation portion of the aircraft 1.

航空機１の発熱部は、例えば、航空機１の電子機器９、バッテリ１０、及びエンジン６の少なくとも一つを含む。また、発熱部は、電子機器９の筐体を含む。航空機１の放熱部は、例えば、航空機１の燃料タンク７を含む。なお、航空機１の放熱部が、航空機１の外部空間（胴体２の外面に面する空間）でもよい。   The heat generating portion of the aircraft 1 includes, for example, at least one of the electronic device 9 of the aircraft 1, the battery 10, and the engine 6. In addition, the heat generating portion includes the housing of the electronic device 9. The heat dissipation unit of the aircraft 1 includes, for example, a fuel tank 7 of the aircraft 1. Note that the heat dissipation unit of the aircraft 1 may be an external space of the aircraft 1 (a space facing the outer surface of the fuselage 2).

複合部材１１は、板状部材である。図４に示す例では、複合部材１１の表面及び裏面のそれぞれに、複合部材２１が配置される。複合部材２１は、炭素繊維で強化されたプラスチックを含む炭素繊維強化プラスチックを有する。   The composite member 11 is a plate-like member. In the example shown in FIG. 4, the composite member 21 is disposed on each of the front and back surfaces of the composite member 11. The composite member 21 has a carbon fiber reinforced plastic including a carbon fiber reinforced plastic.

複合部材２１の炭素繊維強化プラスチックの炭素繊維には、金属が被覆されていない。繊維方向に関する複合部材２１の熱伝導率は、繊維方向に関する複合部材１１の熱伝導率よりも低い。並列方向に関する複合部材２１の熱伝導率は、繊維方向に関する複合部材１１の熱伝導率よりも低い。積層方向に関する複合部材２１の熱伝導率は、繊維方向に関する複合部材１１の熱伝導率よりも低い。   The carbon fiber of the carbon fiber reinforced plastic of the composite member 21 is not coated with metal. The thermal conductivity of the composite member 21 in the fiber direction is lower than the thermal conductivity of the composite member 11 in the fiber direction. The thermal conductivity of the composite members 21 in the parallel direction is lower than the thermal conductivity of the composite members 11 in the fiber direction. The thermal conductivity of the composite member 21 in the stacking direction is lower than the thermal conductivity of the composite member 11 in the fiber direction.

なお、並列方向に関する複合部材２１の熱伝導率は、並列方向に関する複合部材１１の熱伝導率と等しくてもよいし、並列方向に関する複合部材１１よりも低くてもよい。積層方向に関する複合部材２１の熱伝導率は、積層方向に関する複合部材１１の熱伝導率と等しくてもよいし、積層方向に関する複合部材１１よりも低くてもよい。   The thermal conductivity of the composite members 21 in the parallel direction may be equal to the thermal conductivity of the composite members 11 in the parallel direction, or may be lower than that of the composite members 11 in the parallel direction. The thermal conductivity of the composite member 21 in the stacking direction may be equal to the thermal conductivity of the composite member 11 in the stacking direction, or may be lower than that of the composite member 11 in the stacking direction.

金属コート炭素繊維１２の一端部１２Ａ及び他端部１２Ｂのそれぞれに、複合部材２１は配置されない。金属コート炭素繊維１２の一端部１２Ａ及び他端部１２Ｂのそれぞれは、露出する。金属コート炭素繊維１２の一端部１２Ａは、発熱部と接触する。なお、金属コート炭素繊維１２の一端部１２Ａは、発熱部と間隙を介して対向してもよい。金属コート炭素繊維１２の他端部１２Ｂは、放熱部と接触する。金属コート炭素繊維１２の他端部１２Ｂは、放熱部と間隙を介して対向してもよい。   The composite member 21 is not disposed at each of the one end 12A and the other end 12B of the metal-coated carbon fiber 12. Each of the one end 12A and the other end 12B of the metal-coated carbon fiber 12 is exposed. One end 12A of the metal-coated carbon fiber 12 contacts the heating portion. The one end 12A of the metal-coated carbon fiber 12 may be opposed to the heat generating portion via a gap. The other end 12B of the metal-coated carbon fiber 12 is in contact with the heat radiating portion. The other end 12 </ b> B of the metal-coated carbon fiber 12 may be opposed to the heat dissipation portion via a gap.

なお、複合部材２１は、複合部材１１の表面に配置され、複合部材１１の裏面に配置されなくてもよい。複合部材２１は、複合部材１１の裏面に配置され、複合部材１１の表面に配置されなくてもよい。複合部材２１は、複合部材１１の表面及び裏面の両方に配置されなくてもよい。   The composite member 21 may be disposed on the front surface of the composite member 11 and may not be disposed on the back surface of the composite member 11. The composite member 21 may be disposed on the back surface of the composite member 11 and may not be disposed on the surface of the composite member 11. The composite member 21 may not be disposed on both of the front and back surfaces of the composite member 11.

部材２０の繊維方向に金属コート炭素繊維１２が配置されている。部材２０の並列方向及び積層方向のそれぞれにおいて、複数の金属コート炭素繊維１２の間にプラスチック１３が配置されている。繊維方向に関する部材２０の熱伝導率は、並列方向に関する部材２０の熱伝導率及び積層方向に関する部材２０の熱伝導率よりも大きい。   Metal coated carbon fibers 12 are disposed in the fiber direction of the member 20. The plastic 13 is disposed between the plurality of metal-coated carbon fibers 12 in each of the parallel direction and the stacking direction of the members 20. The thermal conductivity of the members 20 in the fiber direction is greater than the thermal conductivity of the members 20 in the parallel direction and the thermal conductivity of the members 20 in the stacking direction.

発熱部の熱は、一端部１２Ａより金属コート炭素繊維１２に吸収される。金属コート炭素繊維１２に吸収された熱は、その金属コート炭素繊維１２を伝わって、他端部１２Ｂより放出（排熱）される。   The heat of the heat generating portion is absorbed by the metal coated carbon fiber 12 from the one end 12A. The heat absorbed by the metal-coated carbon fiber 12 is transmitted along the metal-coated carbon fiber 12 and released (exhaust heat) from the other end 12B.

並列方向に関する部材２０の熱伝導率及び積層方向に関する部材２０の熱伝導率は、繊維方向に関する部材２０の熱伝導率よりも小さい。したがって、金属コート炭素繊維１２の熱は、専ら、繊維方向に移動する。金属コート炭素繊維１２の熱が並列方向及び積層方向に伝達することは抑制される。   The thermal conductivity of the members 20 in the parallel direction and the thermal conductivity of the members 20 in the stacking direction are smaller than the thermal conductivity of the members 20 in the fiber direction. Therefore, the heat of the metal-coated carbon fiber 12 moves exclusively in the fiber direction. Transfer of the heat of the metal-coated carbon fiber 12 in the parallel direction and the stacking direction is suppressed.

また、本実施形態においては、積層方向に関して複合部材１１の表面及び裏面のそれぞれに複合部材２１が配置される。繊維方向、並列方向、及び積層方向のそれぞれの複合部材２１の熱伝達率は、繊維方向に関する複合部材１１の熱伝達率よりも小さい。そのため、金属コート炭素繊維１２の熱が、複合部材２１の表面から放出されることが抑制される。   Further, in the present embodiment, the composite member 21 is disposed on each of the front surface and the back surface of the composite member 11 in the stacking direction. The heat transfer coefficient of each composite member 21 in the fiber direction, the parallel direction, and the stacking direction is smaller than the heat transfer coefficient of the composite member 11 in the fiber direction. Therefore, the heat of the metal-coated carbon fiber 12 is suppressed from being released from the surface of the composite member 21.

以上説明したように、本実施形態によれば、複合部材１１が金属コート炭素繊維１２で強化された金属コート炭素繊維強化プラスチック１４を含み、その金属コート炭素繊維１２の一端部１２Ａが航空機１の発熱部に配置され、金属コート炭素繊維１２の他端部１２Ｂが航空機１の放熱部に配置される。金属コート炭素繊維１２の金属１６は、高い熱伝導率を有する。そのため、発熱部で発生した熱は、金属コート炭素繊維１２を伝わって、効率良く放熱部に放出される。   As described above, according to the present embodiment, the composite member 11 includes the metal-coated carbon fiber reinforced plastic 14 reinforced with the metal-coated carbon fiber 12, and one end 12 A of the metal-coated carbon fiber 12 is the aircraft 1. The other end 12 B of the metal-coated carbon fiber 12 is disposed in the heat dissipation portion of the aircraft 1. The metal 16 of the metal coated carbon fiber 12 has high thermal conductivity. Therefore, the heat generated in the heat generating portion is transmitted to the metal coated carbon fiber 12 and is efficiently released to the heat radiating portion.

また、複合部材１１は、航空機１の強度部材として使用可能である。そのため、従来のような、金属シート又はジャンパー線のような専用の熱流路材を別途設けなくても、発熱部の熱を放熱部に放出できる。そのため、航空機１の重量の増大を抑制しつつ、航空機１の発熱部で発生した熱は、効率良く放熱部に放出される。   The composite member 11 can also be used as a strength member of the aircraft 1. Therefore, the heat of the heat generating portion can be released to the heat radiating portion without separately providing a dedicated heat flow path material such as a conventional metal sheet or a jumper wire. Therefore, while suppressing an increase in the weight of the aircraft 1, the heat generated in the heat generating portion of the aircraft 1 is efficiently released to the heat radiating portion.

本実施形態においては、複合部材１１は、繊維方向と交差する並列方向に配置された複数の金属コート炭素繊維１２を含むプリプレグシート１７を、繊維方向及び並列方向と交差する積層方向に複数積層した積層体を含む。繊維方向に関する部材２０の熱伝導率は、並列方向に関する部材２０の熱伝導率及び積層方向に関する部材の熱伝導率よりも大きい。これにより、熱伝導率に異方性が付与され、発熱部で発生した熱が、並列方向及び積層方向に伝わることが抑制され、放熱部に効率良く放出される。例えば、部材２０の並列方向及び積層方向の少なくとも一方に、加熱したくない部材又は機器が存在する場合、熱伝達率に異方性を持つ部材２０によって、その部材又は機器に熱が伝達されることが抑制される。   In the present embodiment, the composite member 11 is formed by laminating a plurality of prepreg sheets 17 each including a plurality of metal-coated carbon fibers 12 disposed in a parallel direction intersecting the fiber direction in a lamination direction intersecting the fiber direction and the parallel direction Including a laminate. The thermal conductivity of the members 20 in the fiber direction is greater than the thermal conductivity of the members 20 in the parallel direction and the thermal conductivity of the members in the stacking direction. Thereby, anisotropy is imparted to the thermal conductivity, and the heat generated in the heat generating portion is suppressed from being transmitted in the parallel direction and the stacking direction, and is efficiently released to the heat radiating portion. For example, when there is a member or device that does not want to heat in at least one of the parallel direction and the stacking direction of the members 20, heat is transferred to the member or device by the member 20 having anisotropy in heat transfer coefficient Is suppressed.

本実施形態においては、複合部材１１は板状部材であり、部材２０は、複合部材１１の表面及び裏面の一方又は両方に配置された炭素繊維強化プラスチックを含む複合部材２１を備える。これにより、複合部材１１が複合部材２１で支持され、強度が維持される。また、複合部材２１の熱伝導率は、繊維方向に関する複合部材１１の熱伝導率よりも小さい。そのため、部材２０の並列方向及び積層方向の少なくとも一方に、加熱したくない部材又は機器が存在する場合、複合部材２１によって、その部材又は機器に熱が伝達されることが抑制される。   In the present embodiment, the composite member 11 is a plate-like member, and the member 20 includes the composite member 21 including a carbon fiber reinforced plastic disposed on one or both of the front surface and the back surface of the composite member 11. Thereby, the composite member 11 is supported by the composite member 21 and the strength is maintained. Further, the thermal conductivity of the composite member 21 is smaller than the thermal conductivity of the composite member 11 in the fiber direction. Therefore, when there is a member or device that you do not want to heat in at least one of the parallel direction and the stacking direction of the members 20, the composite member 21 suppresses the transfer of heat to the member or device.

また、本実施形態においては、金属コート炭素繊維１２の一端部１２Ａ及び他端部１２Ｂのそれぞれは、複合部材２１などによって覆われてなく、露出する。一端部１２Ａが露出しているので、発熱部で発生した熱は、一端部１２Ａを介して、金属コート炭素繊維１２の金属１６に効率良く吸収される。他端部１２Ｂが露出しているので、発熱部で発生し、金属コート炭素繊維１２の金属１６を移動した熱は、他端部１２Ｂを介して、放熱部に効率良く放出される。このように、本実施形態においては、金属コート炭素繊維１２の一端部１２Ａ及び他端部１２Ｂのそれぞれは露出しているので、発熱部で発生した熱は、放熱部から効率良く放出される。   Further, in the present embodiment, each of the one end 12A and the other end 12B of the metal-coated carbon fiber 12 is not covered by the composite member 21 or the like, and is exposed. Since the one end 12A is exposed, the heat generated in the heat generating portion is efficiently absorbed by the metal 16 of the metal-coated carbon fiber 12 through the one end 12A. Since the other end 12B is exposed, the heat generated in the heat generating part and transferred to the metal 16 of the metal coated carbon fiber 12 is efficiently released to the heat dissipation part via the other end 12B. As described above, in the present embodiment, since each of the one end 12A and the other end 12B of the metal-coated carbon fiber 12 is exposed, the heat generated in the heat generating portion is efficiently released from the heat radiating portion.

本実施形態においては、航空機１の発熱部は、航空機１の電子機器９を含む。航空機１の放熱部は、航空機１の燃料タンク７を含む。これにより、航空機１の高度化により、電子機器９の使用が増大し、電子機器９の発熱量が増大しても、電子機器９で発生した熱は、燃料タンク７に効率良く放出される。   In the present embodiment, the heat generating portion of the aircraft 1 includes the electronic device 9 of the aircraft 1. The heat dissipation unit of the aircraft 1 includes a fuel tank 7 of the aircraft 1. As a result, the use of the electronic device 9 is increased as the aircraft 1 is advanced, and the heat generated by the electronic device 9 is efficiently released to the fuel tank 7 even if the calorific value of the electronic device 9 is increased.

＜第２実施形態＞
第２実施形態について説明する。以下の説明において、上述の実施形態と同一又は同等の構成部分については同一の符号を付し、その説明を簡略又は省略する。 Second Embodiment
The second embodiment will be described. In the following description, the component parts identical or equivalent to those of the above-described embodiment are denoted by the same reference numerals, and the description thereof is simplified or omitted.

図５は、部材２０の使用方法の一例を示す断面図である。図５に示すように、航空機１の発熱部が、部材２０のうち、一端部１２Ａと他端部１２Ｂとの間の中央部に配置されてもよい。すなわち、繊維方向に関して金属コート炭素繊維１２の中央部が発熱部に配置され、金属コート炭素繊維１２の一端部１２Ａ及び他端部１２Ｂが放熱部に配置されてもよい。本実施形態においては、一端部１２Ａ及び他端部１２Ｂのそれぞれが、航空機１の放熱部に配置される。発熱部で発生した熱は、一端部１２Ａ及び他端部１２Ｂのそれぞれから放出される。また、金属コート炭素繊維１２の一端部１２Ａ及び他端部１２Ｂのそれぞれが露出することにより、発熱部で発生した熱は、放熱部から効率良く放出される。   FIG. 5 is a cross-sectional view showing an example of how to use the member 20. As shown in FIG. As shown in FIG. 5, the heat generating portion of the aircraft 1 may be disposed at a central portion of the member 20 between the one end 12A and the other end 12B. That is, the central portion of the metal-coated carbon fiber 12 may be disposed in the heat generating portion in the fiber direction, and the one end 12A and the other end 12B of the metal-coated carbon fiber 12 may be disposed in the heat dissipation portion. In the present embodiment, each of the one end 12A and the other end 12B is disposed in the heat dissipation unit of the aircraft 1. The heat generated in the heat generating portion is released from each of the one end 12A and the other end 12B. Further, when the one end 12A and the other end 12B of the metal-coated carbon fiber 12 are exposed, the heat generated in the heat generating portion is efficiently released from the heat radiating portion.

＜第３実施形態＞
第３実施形態について説明する。以下の説明において、上述の実施形態と同一又は同等の構成部分については同一の符号を付し、その説明を簡略又は省略する。 Third Embodiment
A third embodiment will be described. In the following description, the component parts identical or equivalent to those of the above-described embodiment are denoted by the same reference numerals, and the description thereof is simplified or omitted.

図６は、航空機１の主翼３の平面図を示す。リブライン２９の少なくとも一部に、後述するシアタイ（構造材）３１が配置される。   FIG. 6 shows a plan view of the main wing 3 of the aircraft 1. A shear tie (structural material) 31 described later is disposed on at least a part of the rib line 29.

図７は、外板３０とシアタイ３１との位置関係を概略的に示す図である。シアタイ３１は、ストリンガ、リブ等と外板３０とを結合する部材である。本実施形態において、シアタイ３１は、炭素繊維強化プラスチックを含む複合部材２１と、金属コート炭素繊維強化プラスチック１４を含む複合部材１１とによって形成される。   FIG. 7 schematically shows the positional relationship between the outer plate 30 and the shear tie 31. As shown in FIG. The shear ties 31 are members that connect the stringers, ribs, and the like to the outer plate 30. In the present embodiment, the shear tie 31 is formed by the composite member 21 containing carbon fiber reinforced plastic and the composite member 11 containing metal coated carbon fiber reinforced plastic 14.

外板３０とシアタイ３１とは、ファスナ５１によって固定される。ファスナ５１の先端にカラー（ナット）３７が結合されることによって、外板３０とシアタイ３１とが固定される。カラー３７とシアタイ３１との間には、ワッシャ４１及びスペーサ４２が配置される。   The outer plate 30 and the shear tie 31 are fixed by fasteners 51. The outer plate 30 and the shear tie 31 are fixed by connecting the collar (nut) 37 to the tip of the fastener 51. A washer 41 and a spacer 42 are disposed between the collar 37 and the shear tie 31.

カラー３７、ワッシャ４１、及びスペーサ４２は、キャップ４４によって覆われる。キャップ４４は、シアタイ３１に密着するように配置される。   The collar 37, the washer 41 and the spacer 42 are covered by a cap 44. The cap 44 is disposed in close contact with the shear tie 31.

外板３０は、炭素繊維強化プラスチック層３２と、ガラス繊維強化プラスチック層３４と、銅ペイント層３９と、を含む。   The outer plate 30 includes a carbon fiber reinforced plastic layer 32, a glass fiber reinforced plastic layer 34, and a copper paint layer 39.

図８は、本実施形態に係るシアタイ３１の一例を示す断面図である。シアタイ３１は、炭素繊維強化プラスチックを含む複合部材２１と、金属コート炭素繊維強化プラスチック１４を含む複合部材１１とによって形成される。複合部材１１は、複合部材２１によって挟まれる。複合部材１１は、シアタイ３１の一部に設けられる。図８において、シアタイ３１の下端部に、複合部材１１の金属コート炭素繊維１２の一端部１２Ａが配置される。シアタイ３１の左側の上端部に、複合部材１１の金属コート炭素繊維１２の他端部１２Ｂが配置される。図８に示す例では、シアタイ３１の右側の上端部は、複合部材２１によって形成される。   FIG. 8 is a cross-sectional view showing an example of the shear tie 31 according to the present embodiment. The shear tie 31 is formed of a composite member 21 containing carbon fiber reinforced plastic and a composite member 11 containing metal coated carbon fiber reinforced plastic 14. The composite member 11 is sandwiched by the composite member 21. The composite member 11 is provided on a part of the shear tie 31. In FIG. 8, one end 12 </ b> A of the metal-coated carbon fiber 12 of the composite member 11 is disposed at the lower end of the shear tie 31. The other end 12 B of the metal-coated carbon fiber 12 of the composite member 11 is disposed at the upper end portion on the left side of the shear tie 31. In the example shown in FIG. 8, the upper end on the right side of the shear tie 31 is formed by the composite member 21.

シアタイ３１の下端部に、航空機１の発熱部が配置される。シアタイ３１の左側の上端部に、航空機１の放熱部が配置される。   The heat generating portion of the aircraft 1 is disposed at the lower end portion of the shear tie 31. The heat radiating portion of the aircraft 1 is disposed at the upper end portion on the left side of the shear tie 31.

なお、シアタイ３１の右側の上端部に、複合部材１１の金属コート炭素繊維１２の他端部１２Ｂが配置され、航空機１の放熱部が配置されてもよい。なお、シアタイ３１の右側の上端部及び左側の上端部の両方に、複合部材１１の金属コート炭素繊維１２の他端部１２Ｂが配置され、航空機１の放熱部が配置されてもよい。なお、シアタイ３１の右側の上端部及び左側の上端部の一方又は両方に、複合部材１１の金属コート炭素繊維１２の一端部１２Ａが配置され、航空機１の発熱部が配置されてもよい。シアタイ３１の下端部に、複合部材１１の金属コート炭素繊維１２の他端部１２Ｂが配置され、航空機１の放熱部が配置されてもよい。   The other end 12 B of the metal-coated carbon fiber 12 of the composite member 11 may be disposed at the upper end on the right side of the shear tie 31, and the radiator of the aircraft 1 may be disposed. The other end 12 B of the metal-coated carbon fiber 12 of the composite member 11 may be disposed on both the upper end on the right side and the upper end on the left of the shear tie 31, and the heat dissipation unit of the aircraft 1 may be disposed. One end 12A of the metal-coated carbon fiber 12 of the composite member 11 may be disposed on one or both of the right upper end and the left upper end of the shear tie 31, and the heat generating portion of the aircraft 1 may be disposed. The other end 12 B of the metal-coated carbon fiber 12 of the composite member 11 may be disposed at the lower end of the shear tie 31, and the heat dissipation portion of the aircraft 1 may be disposed.

以上説明したように、複合部材１１及び複合部材２１が曲げられていてもよいし、任意の形状（３次元形状）に加工されてもよい。   As described above, the composite member 11 and the composite member 21 may be bent or may be processed into an arbitrary shape (three-dimensional shape).

なお、上述の各実施形態において、金属コート炭素繊維１２に代えて、又は金属コート炭素繊維１２とともに、ピッチ系炭素繊維が配置されてもよい。ピッチ系炭素繊維は、少なくともＰＡＮ炭素繊維よりも高い熱伝導率を有する熱伝導性炭素繊維である。   In the above-described embodiments, pitch-based carbon fibers may be disposed instead of the metal-coated carbon fibers 12 or together with the metal-coated carbon fibers 12. The pitch-based carbon fiber is a thermally conductive carbon fiber having a thermal conductivity at least higher than that of PAN carbon fiber.

なお、上述の各実施形態においては、部材２０が航空機１の構造部材に使用される例について説明した。部材２０は、人工衛星の構造部材に使用されてもよい。   In each of the above-described embodiments, an example in which the member 20 is used as a structural member of the aircraft 1 has been described. The member 20 may be used for a structural member of a satellite.

なお、上述の各実施形態においては、繊維方向と並列方向と積層方向とは、互いに直交することとした。繊維方向と並列方向とが、例えば８０度以上１００度以下の角度で交差してもよい。繊維方向と積層方向とが、例えば８０度以上１００度以下の角度で交差してもよい。並列方向と積層方向とが、例えば８０度以上１００度以下の角度で交差してもよい。   In each of the above-described embodiments, the fiber direction, the parallel direction, and the stacking direction are orthogonal to each other. The fiber direction and the parallel direction may intersect at an angle of, for example, 80 degrees or more and 100 degrees or less. The fiber direction and the laminating direction may intersect at an angle of, for example, 80 degrees or more and 100 degrees or less. The parallel direction and the stacking direction may intersect at an angle of, for example, 80 degrees or more and 100 degrees or less.

１ 航空機
２ 胴体
３ 主翼
４ 水平尾翼
５ 垂直尾翼
６ エンジン
７ 燃料タンク
８ コックピット
９ 電子機器
１０ バッテリ
１１ 複合部材
１２ 金属コート炭素繊維
１２Ａ 一端部
１２Ｂ 他端部
１３ プラスチック
１４ 金属コート炭素繊維強化プラスチック
１５ 炭素繊維
１６ 金属
１７ プリプレグシート
２０ 部材
２１ 複合部材
２９ リブライン
３０ 外板
３１ シアタイ
３２ 炭素繊維強化プラスチック層
３４ ガラス繊維強化プラスチック層
３７ カラー
３９ 銅ペイント層
４１ ワッシャ
４２ スペーサ
４４ キャップ
５１ ファスナ Reference Signs List 1 aircraft 2 fuselage 3 wing 4 horizontal tail 5 vertical tail 6 engine 7 fuel tank 8 cockpit 9 electronic device 10 battery 11 composite member 12 metal coated carbon fiber 12A one end 12B other end 13 plastic 14 metal coated carbon fiber reinforced plastic 15 carbon Fiber 16 Metal 17 Prepreg sheet 20 Member 21 Composite member 29 Librine 30 Outer plate 31 Shear tie 32 Carbon fiber reinforced plastic layer 34 Glass fiber reinforced plastic layer 37 Color 39 Copper paint layer 41 Washer 42 Spacer 44 Cap 51 Fastener

Claims (4)

金属コート炭素繊維及びピッチ系炭素繊維の一方又は両方を含む熱伝導性炭素繊維で強化されたプラスチックを有し、航空機の強度部材として使用される第１複合部材を備え、
繊維方向に関して前記熱伝導性炭素繊維の一端部が前記航空機の発熱部に配置され、前記熱伝導性炭素繊維の他端部が前記航空機の放熱部に配置され、
前記第１複合部材は、前記繊維方向と交差する並列方向に配置された複数の前記熱伝導性炭素繊維を含むプリプレグシートを前記繊維方向及び前記並列方向と交差する積層方向に複数積層した積層体を含み、
前記繊維方向に関する熱伝導率は、前記並列方向に関する熱伝導率及び前記積層方向に関する熱伝導率よりも大きい、部材。 Comprising a first composite member used as a strength member of an aircraft, comprising a thermally conductive carbon fiber reinforced plastic comprising one or both of metal coated carbon fiber and pitch based carbon fiber,
One end of the thermally conductive carbon fiber is disposed in the heat generating portion of the aircraft in the fiber direction, and the other end of the thermally conductive carbon fiber is disposed in the heat radiating portion of the aircraft ;
The first composite member is a laminate in which a plurality of prepreg sheets including a plurality of the heat conductive carbon fibers disposed in a parallel direction intersecting the fiber direction are laminated in a lamination direction intersecting the fiber direction and the parallel direction Including
The thermal conductivity in the fiber direction is greater than the thermal conductivity in the parallel direction and the thermal conductivity in the stacking direction . 金属コート炭素繊維及びピッチ系炭素繊維の一方又は両方を含む熱伝導性炭素繊維で強化されたプラスチックを有し、航空機の強度部材として使用される第１複合部材を備え、
繊維方向に関して前記熱伝導性炭素繊維の中央部が前記航空機の発熱部に配置され、前記熱伝導性炭素繊維の一端部及び他端部が前記航空機の放熱部に配置され、
前記第１複合部材は、前記繊維方向と交差する並列方向に配置された複数の前記熱伝導性炭素繊維を含むプリプレグシートを前記繊維方向及び前記並列方向と交差する積層方向に複数積層した積層体を含み、
前記繊維方向に関する熱伝導率は、前記並列方向に関する熱伝導率及び前記積層方向に関する熱伝導率よりも大きい、部材。 Comprising a first composite member used as a strength member of an aircraft, comprising a thermally conductive carbon fiber reinforced plastic comprising one or both of metal coated carbon fiber and pitch based carbon fiber,
The central portion of the thermally conductive carbon fiber in the fiber direction is disposed in the heat generating portion of the aircraft, and one end and the other end of the thermally conductive carbon fiber are disposed in the heat radiating portion of the aircraft ;
The first composite member is a laminate in which a plurality of prepreg sheets including a plurality of the heat conductive carbon fibers disposed in a parallel direction intersecting the fiber direction are laminated in a lamination direction intersecting the fiber direction and the parallel direction Including
The thermal conductivity in the fiber direction is greater than the thermal conductivity in the parallel direction and the thermal conductivity in the stacking direction . 前記第１複合部材は板状部材であり、
前記第１複合部材の表面及び裏面の一方又は両方に配置され、炭素繊維で強化されたプラスチックを含み、前記第１複合部材よりも熱伝導率が低い第２複合部材を備え、
前記熱伝導性炭素繊維の一端部及び他端部のそれぞれは露出する請求項１又は請求項２に記載の部材。 The first composite member is a plate-like member,
The first is arranged on one or both of the front and rear surfaces of the composite member, seen including a plastic reinforced with carbon fibers, a second composite member lower thermal conductivity than the first composite member,
The member according to claim 1 or 2 , wherein each of the one end portion and the other end portion of the heat conductive carbon fiber is exposed. 前記発熱部は、前記航空機の電子機器を含み、前記放熱部は、前記航空機の燃料タンクを含む請求項１から請求項３のいずれか一項に記載の部材。 The member according to any one of claims 1 to 3 , wherein the heat generating portion includes an electronic device of the aircraft, and the heat radiating portion includes a fuel tank of the aircraft.
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